Small arrayed swirler system for reduced emissions and noise

ABSTRACT

A combustor assembly employs an ejector assembly having an array of swirlers to influence the flow field, reduce unburned fuel emissions, enhance flame stabilization and reduce noise within a combustion chamber. The swirlers are located downstream of a premixing section and are designed to correct bias of fuel-air mixture residence times in the combustion chamber. The ejector assembly may be tuned to any particular configuration by varying the swirl number in order to produce recirculation zones that ensure out of phase nature of the heat release within the combustion chamber.

FIELD OF TECHNOLOGY

The present disclosure relates to gas turbine engines, and more particularly, to an improved combustor assembly employing an array of swirlers to influence the flow field and enhance flame stabilization within a combustion chamber.

BACKGROUND

Gas turbine engines are known to include a compressor for compressing air, a combustor for producing a hot gas by burning fuel in the presence of the compressed air produced by the compressor, and a turbine for expanding the hot gas to extract power. Gas turbine engines using annular combustion systems typically include a plurality of individual burners disposed in a ring about an axial centerline for providing a mixture of fuel and air to an annular combustion chamber disposed upstream of the annular turbine inlet vanes. Other gas turbines use can-annular combustors wherein individual burner cans feed hot combustion gas into respective individual portions of the arc of the turbine inlet vanes. Each can includes a plurality of main burners disposed in a ring around a central pilot burner.

During operation, the combustion flame can generate combustion oscillations, also known as combustion dynamics. Combustion oscillations in general are acoustic oscillations which are excited by the combustion itself. The frequency of the combustion oscillations may be influenced by an interaction of the combustion flame with the structure surrounding the combustion flame. Since the structure of the combustor surrounding the combustion flame is often complicated, and varies from one combustor to another, and because the combustion flame itself may vary over time, it is difficult to predict the frequency at which combustion oscillations occur. As a result, combustion oscillations may be monitored during operation and parameters may be adjusted in order to influence the interaction of the combustion flame with its environment.

A combustion flame emits sound energy during combustion. A more uniform flame will generate more uniform acoustics, but perhaps with higher peak amplitude at a particular frequency than a less uniform flame. When an emitted frequency of combustion coincides with a resonant frequency of the combustion chamber the system may operate in resonance, and the resulting combustion dynamics may damage the gas turbine components, or at least reduce their lifespan.

One known way to reduce the interaction of the combustion flame with the combustion acoustics is to reduce the coherence of the flame, i.e. reduce the spatio-temporal uniformity of the flame. A flame with less uniform combustion throughout its volume is likely to perturb the gas turbine less than a uniform flame because the energy released is spatially distributed and therefore decreases its coupling to the system resonant frequencies or acoustic modes. As a result, combustion dynamics of flames with less uniform combustion throughout its volume are less likely to be exacerbated than by a more uniform flame. Creating a less uniform fuel-air mixture would be helpful.

During the combustion of gas, pollutants such as, but not limited to, carbon monoxide (“CO₂”), unburned hydrocarbons (“UHC”), and nitrogen oxides (“NO_(x)”) may be formed and emitted into an ambient atmosphere. Because of stringent emission control standards, it is desirable to control emissions of such pollutants by the suppressing formation of such emissions. It would be helpful to reduce such emissions to target levels so as to meet government emission levels. This may be accomplished by minimizing the resident time unburnt fuel resides within the combustion chamber.

In the case of industrial gas turbines which primarily burn gaseous fuels (e.g. for power generation purposes) several contradictory requirements with respect to premixer, flame stabilizer and combustion chamber design for lean partially-premixed systems require an achievement of tradeoffs between cost, complexity, robustness to operating conditions and lowered emissions. For example, a perfectly premixed system is more prone to thermo-acoustic oscillations leading to hardware damage while a less uniformly mixed system results in increase of pollutant formation, in particular, NOx. Another issue is that the orientation of the combustion chamber within any engine architecture can impose constraints on burner geometry, that result in unsymmetrical stream tubes within the combustion chamber, which in turn, lead to a variation of mixture residence times leading to lowered combustion efficiencies, increased risk of undesirable events such as poor flame stabilization, flashback, blow out and less than adequate mixing of post-flame gases which is required for the fast burnout of Carbon Monoxide to Carbon Dioxide. The time taken for the latter is a hard constraint in many systems as it is obtained—in a simple example—as the time taken for the mixture to travel from the flame front to the chamber exhaust, which is determined by the engine shaft length. However, overly long residence times (desirable for oxidation of CO) can increase the amount of NOx formation at high temperatures.

For this reason, it is helpful that the flame stabilizer to include these considerations into its design and be able to function in a robust manner in engine architectures of reasonable variation. Typically, symmetric systems are the goal for reasons of ease of operation and lowered costs but rarely achieved in real-life. Still another tradeoff lies in ensuring that the mixture in the premixer is expelled before it auto-ignites. At typical engine operating conditions, the compressed air can reach over 800K and pressures in excess of 500 psi. At these conditions, simple fuels like CH4 can ignite within 40 milliseconds while more reactive fuels such as diesel can ignite in less than 2 milliseconds. Premixer design should ensure that adequate uniformity of mixing is achieved within a timescale that is lower than the ignition delay of the range of fuels required. With safety factors that can range from a tenth to twentieth of the timescales, it is difficult to achieve a uniformity of mixture unless the premixer design employs high turbulence intensities.

Fuel injection into a continuous burning combustion chamber as, for example, in a gas turbine engine has posed continuing design problems. Difficulties have been encountered in injecting fuel in a highly dispersed manner so as to achieve complete and efficient combustion of the fuel, and at the same time minimize the occurrence of fuel-rich pockets which upon combustion produce carbon, smoke, or unburned hydrocarbon pollutants. Fuel injection difficulties have been further complicated by the introduction of gas turbine engines having increased combustor pressure levels. Existing fuel spray atomizer efficiency decreases as combustor pressure is increased, resulting in a more non-uniform dispersion of fuel, together with an increase in the fuel-rich zones within the combustion chamber which cause reduced burner efficiency, excessive exhaust smoke, and a non-uniform heating of the combustor shell, a condition commonly referred to as hot streaking, which can lead to rapid deterioration of the shell.

High fuel pressure spray atomizers also have not proved entirely satisfactory because of the present limitations on fuel pump pressure. Systems for vaporizing fuel upon injection into the combustor have also proved to be severely limited due to the dependence of the vaporization process on the temperature of the fuel and air entering the combustor.

In view of the aforementioned challenges, there is a need to provide an improved combustor assembly and method of operation that provides improved flame stabilization for reducing emissions and noise. The system may be tunable and include a plurality of swirlers that results in a heat release that is designed to be out of phase with each other so as to ensure that thermo-acoustic couplings are dampened.

BRIEF DESCRIPTION OF THE DRAWINGS

While the claims are not limited to a specific illustration, an appreciation of the various aspects is best gained through a discussion of various examples thereof. Referring now to the drawings, exemplary illustrations are shown in detail. Although the drawings represent the illustrations, the drawings are not necessarily to scale and certain features may be exaggerated to better illustrate and explain an innovative aspect of an example. Further, the exemplary illustrations described herein are not intended to be exhaustive or otherwise limiting or restricted to the precise form and configuration shown in the drawings and disclosed in the following detailed description. Exemplary illustrations are described in detail by referring to the drawings as follows:

FIG. 1 illustrates a schematic view of a gas turbine engine employing the improvements discussed herein;

FIG. 2 illustrates a partial perspective view of combustor for a gas turbine engine employing an exemplary swirler assembly for use with a combustor;

FIG. 3 illustrates a perspective view of a swirler ring assembly having a plurality of swirlers disposed circumferentially around a ring structure;

FIG. 4 is a side elevational view of the FIG. 3 swirler ring assembly;

FIG. 5 is a front elevational view of the FIG. 3 swirler ring assembly;

FIG. 6 is a front elevational view of an alternative swirler ring assembly, employing an array of swirlers and slots;

FIG. 7 is a front elevational view of another alternative swirler ring assembly, employing ports located adjacent to the swirlers; and

FIG. 8 is a schematic view of a section of a combustor, showing a primary section and a secondary section, and fluid flow within the combustion chamber.

DETAILED DESCRIPTION

An exemplary embodiment discloses an improved combustor assembly and method that overcomes traditional combustor challenges by employing a tunable flame stabilizer assembly in the injector section of an engine, such as, but not limited to, a gas turbine. The stabilizer assembly includes an array of swirlers located downstream of a premixing section that are designed to correct bias of fuel residence times in the combustion chamber. The flame stabilizer could employ an ejector like ring having a plurality of circumferentially spaced swirlers that induce a premixed air/fuel mixture to the combustion chamber. The stabilizer may be tuned by varying the swirl number in order to produce recirculation zones that ensure out of phase nature of the heat release within the combustion chamber. Individual premixing channels receive fuel and pre-heated air which rely on turbulent mixing set up by vortices generated within the premixer channel and expels the mixed material through each swirler and into the combustion chamber where it is burned to provide energy for the turbine section.

An alternative embodiment provides a flame stabilizer assembly featuring a hybrid arrangement of swirlers and jet nozzles for passing the air/fuel mixture through an injector. This allows a compensation of more drastic combustion chamber residence time biases. Another embodiment features a premixer that injects non-uniform fuel flow through a series of ports that are located near the array of swirlers. It will be appreciated that a combination of the embodiments disclosed herein may be employed so as to include a variety of swirler, slot and injector configurations so as to accommodate a wide variety of combustion chambers typical of gas turbine engines

A method of operating an injector for a combustor includes mixing air and fuel into a first primary mixing chamber, introducing the air/fuel mixture into a plurality of swirlers, directing the mixture into a combustion chamber, and then igniting the mixture. The method further includes mixing air and fuel into a secondary mixing chamber, introducing the mixed air and fuel into a common secondary swirler, directing the mixture into a secondary combustion zone within the combustion chamber, and igniting the mixture. The method may further include additional primary mixing chambers that are circumferentially spaced around the central axis 28 of the machine 10. Likewise, each primary mixing chamber has an associated swirler that is associated with it which operates in a manner similar to that discussed for the first primary mixing chamber.

The resulting combustor assembly is an array of swirlers that generates a series of small recirculation zones in order to reduce the length of scales of coherent flame structures. The number of swirlers can be varied in order to produce recirculation zones that ensure out of phase nature of the heat released within the combustor. The swirl number is tuned to the asymmetry of the combustion chamber so that swirlers with a low swirl number are situated at locations which give rise to stream tubes of high residence time. By contrast, swirlers with a high swirl number are situated at locations which rise to stream tubes of low residence time. As such, the swirler may be used to correct bias of residence times in the combustion chamber. It will be appreciated that the number of swirlers may range from 0 to 18 (as shown), or more.

FIG. 1 illustrates a gas turbine engine 10, which includes a fan 12, a low pressure compressor and a high pressure compressor, 14 and 16, a combustor 18, and a high pressure turbine and low pressure turbine, 20 and 22, respectively. The high pressure compressor 16 is connected to a first rotor shaft 24 while the low pressure compressor 14 is connected to a second rotor shaft 26. The shafts extend axially and are parallel to a longitudinal center line axis 28.

Ambient air 30 enters the fan 12 and is directed across a fan rotor 32 in an annular duct 34, which in part is circumscribed by fan case 36. The bypass airflow 38 provides engine thrust while the primary gas stream 40 is directed to the combustor 18 and the high pressure turbine 20. The gas turbine engine 10 includes an improved combustor 18 having a unique flame stabilizer assembly 42 for improved heat release in the combustion chamber.

FIG. 2 illustrates a perspective view of the inside of a combustor assembly 18 showing some of the components of the flame stabilizer assembly 42. It will be appreciated that the outer liner wall has been removed so as to provide improved understanding of the components of the stabilizer assembly 42.

The stabilizer assembly includes an injector ring 44, a primary mixing chamber or duct 46, a secondary mixing chamber or duct 48, a secondary swirler 50, a plurality of primary swirlers 52, and a liner 54. A primary combustion zone 56 connotes the area in which the air/fuel mixture from an array primary swirlers 52 deposit atomized fuel particles or fuel-air mixtures that are ready to be combusted. A secondary combustion zone 58 connotes an area where the secondary swirler 50 deposits atomized fuel particles or fuel-air mixtures that are ready to be combusted.

The secondary mixing chamber 48 receives air flow from air duct 74 and fuel flow from fuel supply 64 in which the air and fuel are mixed and exited at outlet 66. Vanes 68 force the mixture from chamber 48 to exit through the outlet 66, thereby directing the heated, pressurized fuel/air mixture or atomized fuel/air mixture to be stabilized prior to combustion.

The inner liner 54 is made of conventional material and extends between the secondary swirler 50 and the array of swirlers 52. The liner 54 extends axially from the injector ring 44 and terminates near a flange 70 that is formed near one end of a partition member 72. An air duct 74 supplies cooling air to the inner liner 54 where ports 76 deliver the air to the primary combustion zone 56. The secondary duct 48 receives secondary fuel 78 and secondary air 80 that collectively form the fuel/air mixture 64. See FIG. 8.

With reference to FIGS. 2 and 3, the injector ring 44 can be, but is not limited to, an annularly shaped member that circumscribes the inner liner 54. The ejector 44 has an inner diameter 82 and an outer diameter 84 that are concentric with axis 28 and is uniform in width and depth but variations thereof to accommodate non-circular jets are included as well. A plurality of passages 86 extend from the face 88, through the body 90 of the ejector 44, and extends to back 92 of the injector 44. Each passage 86 receives a swirler 52 for advancing the mixed air/fuel 64′ that is generated from the primary duct 46.

FIG. 4 illustrates a side view of the injector ring 44 whereby the face 88 has an angle alpha that forms a relief from outer surface 94. This configuration differs from the flat face 88 that is shown in the embodiment of FIG. 2 whereby the flat face 88 is basically normal to the axis 28 of the machine. However, by locating the face 88 at an angle alpha as is shown in FIG. 4, each passage 86 is oriented to no longer be normal to the axis 28, but instead is redirected to provide a path extending a flow of atomized fuel or fuel-air mixture into the combustion chamber 18 that is non-normal. This causes small pockets of recirculation zones that bleed over to adjacent recirculation zones from the adjacent swirler 52. The angle alpha may be a variety of configurations including, but not limited to 21, 24 or 26 degrees.

The injector ring 44 further has a first annular grove 96 and a second annular groove 98. The grooves extend around the periphery of the ring 44 and form a fluid flow path. The ring 44 is made of suitable material to withstand the temperatures that are traditionally present in the combustor applications.

FIG. 5 illustrates the injector ring 44 from the front elevational view. The ring 44 is shown in this exemplary embodiment having 18 swirlers 52 that are spaced apart around the circumference of the body 90. Each swirler is positioned within a passage 86 for providing its own small recirculation zone within the primary combustion zone 56. It will be appreciated that more or less swirlers may be provided. While 5 blades 100 are shown on each swirler 52, it will be appreciated that the number of blades 100 on each swirler 52 may vary. Thus, the injector ring 44 may be tuned by varying the number of swirlers 52 and blades 100 that are employed. Also, the alpha angle may be varied so as to change the pitch or trajectory 258 (FIG. 8) of delivery of atomized fuel air mixture into the primary combustion zone 56.

FIG. 6 illustrates an alternative embodiment of an injector ring 150 having a blend of swirlers 52 and slots 152. The swirler designs are similar to those mentioned above, and in this variation, 9 swirlers are presented in an array between the 6 o'clock to 12 o'clock positions. By contrast, an array of slots 152 are aligned between the 1 o'clock and 5 o'clock positions. It will be appreciated that this mix of swirlers and slots may be positioned at other locations and they could even be mixed intermittently between one another. The slots 152 are shown as arcuate shaped rectangles and they have a space 154 of material separating each adjacent slot 152. The slots 152 are ports or jets that extend through the body 90 and provide a passageway for the jet flames to enter the primary combustion zone 56. The slots 152 may be fed with a fuel/air mixture at a different equivalence ratio compared to the swirlers 52. The recirculation zone 202 (FIG. 8) may be designed to have preferred performance characteristics that are based upon a combination of swirlers 52 and slots 152.

FIG. 7 depicts another alternative embodiment injector ring 200 that is designed to inject a non-uniform fuel flow into the primary combustor zone 56. This intentional mixture of non-uniformity however is with small acceptable margins of non-uniformity so as to contain the NOx emissions within industry standards. This alternative embodiment 200 is a variant of the FIG. 5 assembly, whereby jets 204 and 206 are positioned adjacent to swirlers 52 at positions around the circumference of the body 90. The jets 204 and 206 are shown in pairs in alternating patterns spaced adjacent to a pair of swirlers 52. It will be appreciated that the jets may be positioned in between each swirler 52. The jet 204 is shown larger in diameter than jet 206 and they could be reversed to have the smaller jet on the outside of the ring 200 with the larger jet 204 positioned near the inside diameter of the ring 200.

FIG. 8 illustrates a schematic fluid diagram 250 of the fuel/air mixtures entering the combustion zones within the combustor 18. The primary mixer or duct 46 represents the fluid mixing channel where primary fuel 252 and primary air 254 are mixed to form the fuel/air mixture 64′. The mixed fuel 64′ is delivered to an injector ring of the styles shown at ejectors 44, 150, or 200. The mixture 64′ is atomized in the process and passes through the injector ring 44 to the primary combustion zone 56 which is at a leading portion of the combustor 18.

The example shown has the injector outlet passage 86 configured at an angle alpha which results in the plume 256 of atomized fuel or fuel-air mixture to enter the primary combustion zone 56 along a trajectory 258. This results in small pockets 260 of recirculation zones that are associated with each swirler 52. Thus, for each swirler 52 that is spaced around the periphery of the injector ring 44, a small pocket 260 of recirculation zone is induced into the primary combustion zone 56. The small pockets 260 overlap with one another so as to enhance combustion and to maintain ignition within the combustor 18. The performance of each small pocket 260 is controlled by, in part, the angle alpha, the number and design of each blade for each swirler 52, and other controllable features. As such the ejector ring 44 is tunable to afford different performance characteristics which results in different fluid flow patterns within the primary combustion zones 56 and the secondary combustion zone 58.

The secondary duct 48 is a mixing channel that receives the secondary fuel 78 and the secondary air 80 which in turn is heated and mixed to form a mixture 64 within the flow path 60. The mixture 64 is delivered to the jet or swirler 50 and exits into the secondary combustion zone 58 and forms a recirculation pattern 262. The recirculation patterns 262 and 202 may combine within the trailing portion of the combustor 18 so as to enhance ignition of the fuel particles and create heat for rotating the turbines.

The process of operating the engine 10 using the injector ring 44 will now be presented with reference to schematic of FIG. 8. Fuel and air mixtures are induced to the primary duct 46 and secondary duct 48. The fuel mixtures 64 and 64′ advance to their associated jets, 50 and 52, which in turn delivers atomized fuel particles to their respective combustion zones 58 and 56. Ignition now occurs within the combustion zones which produced heat for driving the turbines 20 and 22.

The method of operation may be tuned by changing the injector ring 44 to have a variety of swirlers 52 and slots 152 and/or jets 204 and 206, or any combination thereof. By modifying this arrangement of components, the recirculation zone characteristics 202 within the primary combustion zone 56 can be tuned to a desired performance. This results in part due to the individual recirculation zones 260 that are generated by each individual swirler 52, slot 152 and or jets 204 and 206. Thus by locating a swirler 52, slot 152 and/or jet 204, 206 at predetermined locations around the ejector ring 44, the user can influence the combustion, emission, and energy generated by the engine 10.

It will be appreciated that the aforementioned method and devices may be modified to have some components and steps removed, or may have additional components and steps added, all of which are deemed to be within the spirit of the present disclosure. Even though the present disclosure has been described in detail with reference to specific embodiments, it will be appreciated that the various modifications and changes can be made to these embodiments without departing from the scope of the present disclosure as set forth in the claims. The specification and the drawings are to be regarded as an illustrative thought instead of merely restrictive thought.

All terms used in the claims are intended to be given their broadest reasonable constructions and their ordinary meanings as understood by those knowledgeable in the technologies described herein unless an explicit indication to the contrary is made herein. In particular, use of the singular articles such as “a,” “the,” “said,” etc. should be read to recite one or more of the indicated elements unless a claim recites an explicit limitation to the contrary. 

What is claimed is:
 1. A combustion apparatus comprising: a combustor liner; an injector assembly including a primary combustion system and a secondary combustion system; and a combustion chamber having an axis extending therethrough.
 2. The combustion apparatus according to claim 1, further comprising a gas turbine engine.
 3. The combustion apparatus according to claim 1, wherein the primary combustion system includes a mixing chamber, and a nozzle, the mixing chamber has ports for receiving air and fuel, the mixing chamber further has an exit located adjacent to the nozzle.
 4. The combustion apparatus according to claim 1, wherein the primary combustion system includes an injector ring, the injector ring is spaced apart from the combustor liner.
 5. The combustion apparatus according to claim 1, wherein the primary combustion system includes an injector member, the injector member has an inside diameter, an outside diameter, a body, a front face and a rear face, the front face has a plurality of passage openings, the passage openings extend through the body.
 6. The combustion apparatus according to claim 5, further comprising a swirler located within at least one passage opening.
 7. The combustion apparatus according to claim 5, further comprising a plurality of swirlers located around a circumference of the body of the injector member.
 8. The combustion apparatus according to claim 5, wherein the front face is not normal to the axis that extends through the combustor.
 9. The combustion apparatus according to claim 5, further comprising a swirler located at one portion of the injector member, and at least one slot located at another portion of the injector member.
 10. The combustion apparatus according to claim 5, further comprising an array of swirlers that are spaced apart from one another around at least a portion of the injector member.
 11. The combustion apparatus according to claim 5, wherein the injector member is ring shaped and has an array of spaced apart swirlers that are located downstream from a mixing channel.
 12. The combustion apparatus according to claim 1, wherein the secondary combustion system includes a mixing channel, a swirler, and a passageway in communication with the mixing channel.
 13. The combustion apparatus according to claim 1, wherein the secondary combustion system has a flow path that delivers combustible material to a secondary combustion zone, and the primary combustion system has a flow path that delivers combustible material to a primary combustion zone.
 14. A combustion member comprising: a combustor liner; an injector assembly, the injector assembly includes a body having a plurality of spaced apart apertures circumferentially located around the body, a swirler is located within at least one aperture; a primary combustion system, the primary combustion system includes a mixing space, the mixing space includes openings for receiving fuel and air; and a secondary combustion system, the secondary combustion system includes a mixing space that has openings for receiving fuel and air.
 15. The combustion apparatus according to claim 14, wherein the injector assembly further includes an annular channel that is spaced upstream of the swirler.
 16. The combustion apparatus according to claim 14, further including a swirler located within each aperture that is located around the ring.
 17. The combustion apparatus according to claim 14, wherein the spaced apart apertures include a combination of slots and circular shaped openings, the swirler is located within the circular shaped opening, and the slot performs the function as a jet.
 18. The combustion apparatus according to claim 14, wherein the spaced apart further includes a swirler for introducing fuel particles or fuel-air mixtures to a secondary combustion zone that is located within a combustor.
 19. The combustion apparatus according to claim 14, further comprising a combustor shell.
 20. A combustion member comprising: a combustor liner, the liner has ports for facilitating the entry of air; an injector assembly, the injector assembly includes an annular shaped ring having a plurality of spaced apart apertures circumferentially located around the ring, a swirler is located within each aperture; a primary combustion system, the primary combustion system includes a mixing space, the mixing space includes openings for receiving fuel and air; a primary combustion zone located downstream of the swirler; a secondary combustion system, the secondary combustion system includes a second swirler and a mixing space that has openings for receiving fuel and air; and a secondary combustion zone located downstream of the second swirler. 